JP2007218255A - Double turbofan engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a supersonic type aircraft turbofan engine improved in variable cycle performance, efficiency, and noise damping performance. <P>SOLUTION: The turbofan engine 10 includes first and second fans 14, 16 each independently connected to turbines 26, 24. A first bypass pipe passage 46 is arranged to surround a core engine and communicate with the second fan 16. A second bypass pipe passage 48 surrounds the first bypass pipe passage 46 and is communicated with the first fan 14. A first discharge nozzle 66 is connected to both the core engine and the first bypass pipe passage 46. Further, a second discharge nozzle 68 is connected to the second bypass pipe passage 48. <P>COPYRIGHT: (C)2007,JPO&INPIT

Description

  The present invention relates generally to gas turbine engines, and more particularly to variable cycle engines for powering aircraft flying at supersonic speeds.

  A typical aircraft turbofan gas turbine engine includes a single stage fan driven by a low pressure turbine (LPT). The fan is followed by a multistage axial compressor that further compresses the air mixed with fuel in the combustor to generate hot combustion gases. Energy is extracted from the combustion gases in the high pressure turbine (HPT) that powers the compressor.

  The fan and compressor are connected to the corresponding rotors of LPT and HPT by independent drive shafts or spools. In this way, the fan and compressor operating lines can be controlled independently during various periods of the flight envelope, including takeoff, ascent, level flight, landing approach, and landing on the runway.

  The turbofan engine is arranged in two separate configurations. One configuration includes a short conduit or nacelle that surrounds the fan in a high bypass ratio configuration, which has separate fans and core engine exhaust nozzles, with the air compressed by the fan and the core engine. The generated combustion gas is released separately.

  The second configuration of the turbofan engine includes a long line or nacelle that surrounds the fan and extends to the rear end of the engine in a common exhaust nozzle, which exhaust nozzles contain compressed fan air and combustion exhaust gas. Release both.

  In both configurations, both the short and long bypass lines surround the core engine to bypass or bypass a portion of the compressed fan air around the core engine. The core engine includes a high pressure compressor having limited flow capacity therein.

  In a short nacelle configuration, the fan bypass line is correspondingly short and terminates in a separate fan nozzle.

  In long line configurations, the bypass line extends from the fan downstream of the LPT and typically recombines the bypass air into the combustion exhaust stream before releasing it to a common exhaust nozzle.

  A common turbofan aircraft engine and its two independent rotors are typically configured to power aircraft at subsonic speeds well below Mach 1.

  However, in supersonic military aircraft or civilian aircraft, it is sought to increase the amount of propulsion thrust required to accelerate the aircraft to a supersonic speed faster than Mach 1 and maintain that supersonic speed during long horizontal flight operations. This substantially increases the size, weight, and complexity of the turbofan engine. Supersonic business jets (SSBJs) designed for sustained supersonic level flight operation also require commercially viable engine efficiency and exhaust noise tuning tolerance levels.

  Noise generation in supersonic aircraft is an important design issue to address various government noise regulations (usually the most severe in the immediate vicinity of an airport).

  Therefore, the prior art is rich in various configurations of variable cycle turbofan engines that are configured to power aircraft, particularly at supersonic speeds. The size, weight, and complexity of these various variable cycle turbofan engines vary dramatically with their aerodynamic efficiency and the noise level generated during operation. In designing the various parts of a supersonic aircraft engine, significant compromises must be made to balance conflicting design objectives in order to achieve high performance.

  One form of variable cycle engine includes FLADE (an acronym for “fan on blade”). FLADE is a special form of fan that includes relatively large fan blades that have a radially outer tip extension that is limited by a partial span integral side plate. The FLADE airfoil, or the outer portion of the fan blades on the side plate, is specially configured so that the aerodynamic cross section effectively compresses the tip air flowing downstream in the corresponding annular bypass line surrounding the core engine Has been. In that case, this FLADE bypass air can use a special exhaust nozzle in various forms to reduce acoustic noise during the desired period of the flight envelope.

  An important issue when incorporating FLADE into a turbofan engine is that it generates additional centrifugal forces during operation that must be received by the inner airfoil and the supporting rotor disk. is there. The outer FLADE airfoil and its integral inner side plate produce a large centrifugal load during the rotational operation of the fan. Therefore, in order to carry the centrifugal load within the range of a suitable stress limit and to ensure the life of the fan, it is necessary to increase the thickness of the inner rotor foil and the supporting rotor disk.

  Increasing the fan airfoil reduces the aerodynamic efficiency and performance of the airfoil, which correspondingly reduces the overall efficiency of the engine.

  Thus, FLADE can be used to provide compressed air with an increased fan pressure ratio for acoustic nozzles in a turbofan engine with the same noise level as a larger, lower fan pressure ratio engine cycle. The FLADE mixed-flow turbofan engine can exhibit performance improvements over the FLADE variable-cycle engine in subsonic horizontal flight configurations (but superior to the conventional mixed-flow turbofan engine) Is very small).

  Engines with FLADE can benefit from increased thrust per unit of airflow at the expense of increased centrifugal load from the FLADE airfoil and a corresponding increase in the weight of the engine. Yes, as well as the aerodynamic performance penalty due to the thicker support fan airfoil under FLADE.

  In addition, the introduction of FLADE into turbofan engines typically includes an inlet guide vane (IGV) before the fan stage with FLADE, as well as an outlet guide vane (OGV) after the stage with FLADE. These guide vanes are used to increase aerodynamic efficiency, but the length of the engine must be correspondingly increased, and the weight and complexity must be increased accordingly.

Next, when constructing a practical supersonic aircraft engine, the dilemma faced by engine designers is usually the aerodynamic configuration, mechanical strength, exhaust noise of various parts of the interrelated turbofan engine, A delicate balance between size, weight, and complexity.
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  Accordingly, it is desirable to provide a supersonic aircraft turbofan engine with improved performance and efficiency, and noise attenuation.

  The variable cycle turbofan engine includes first and second fans that are independently coupled to the turbine. The first bypass pipe is arranged so as to surround the core engine and communicate with the second fan. The second bypass conduit surrounds the first bypass conduit and communicates with the first fan. The first exhaust nozzle is connected to both the core engine and the first bypass line. The second exhaust nozzle is connected to the second bypass conduit.

  The preferred exemplary embodiments, together with further objects and advantages thereof, will be more particularly described in the following detailed description in conjunction with the accompanying drawings.

  Illustrated schematically in FIG. 1 is a variable cycle turbofan gas turbine engine 10 configured to power a flying aircraft (not shown) at a flight speed in excess of Mach 1 from subsonic to supersonic. It is. The turbofan engine 10 is axisymmetric about the longitudinal or axial central axis 12 and can be suitably mounted on an aircraft wing or body.

  The engine 10 includes a first fan 14, a second fan 16, a high pressure compressor 18, a combustor 20, a first or high pressure turbine (HPT) 22, a second or medium pressure turbine (IPT) 24, A third or low-pressure turbine (LPT) 26 and an exhaust line 28 are disposed that are coaxially connected in series along the central axis 12.

  Ambient air 30 enters the engine through its intake and is subsequently compressed by a fan and compressor and is mixed with fuel in the combustor 20 to generate hot combustion gases 32. Within the three turbines, energy is extracted from the combustion gases, powering the fan and compressor, and the combustion gases are discharged from the exhaust line 28.

  The first fan 14 is connected to a first spool or drive shaft 34. The second fan 16 is connected to the second turbine 24 by a second spool or drive shaft 36. The compressor 18 is connected to the first turbine 22 by a third spool or a drive shaft 38. These three drive shafts are coaxial with each other and concentric.

  The first annular casing 40 surrounds the core engine including the compressor 18, the combustor 20, and the HPT 22, and extends rearward beyond the second turbine 24 and the third turbine 26. The second annular casing 42 is spaced outward from the first casing 40 in the radial direction or concentrically with the outer side. The third annular casing 44 is concentrically spaced from the second casing 42 outward in the radial direction.

  The first casing 40 and the second casing 42 define an annular inner or first bypass line 46 therebetween in the radial direction, the first bypass line connecting the compressor 18 in communication with the second fan 16. The core engine is coaxially enclosed, and a part of the compressed air 30 is received from the second fan. The first bypass conduit 46 extends in the axial direction around the core engine in a length portion from the front inlet end portion behind the second fan 16, and the rear end portion is common to the core engine. It terminates in the exhaust line 28. Therefore, the core engine is bypassed.

  The second casing 42 and the third casing 44 define an annular outer or second bypass line 48 therebetween in the radial direction, the second bypass line being connected to the second fan 16 and the inner bypass line 46. Are communicated with the radially outer front end of the first fan 14. Outer bypass line 48 is located at its outlet end located at the rear end of inner bypass line 46 axially behind core engine and third turbine 26 from its inlet end immediately after first fan 14. In the length part up to, it extends in the axial direction.

  These bypass lines 46, 48 are concentric with each other and extend from the upstream fan for the majority of the axial length of the engine by the length of the line configuration, and the inner core engine is separated from the two fans. Bypass with two concentric airflows.

  Specifically, the first fan 14 has a large diameter, extends radially outward beyond the radial span of the small diameter second fan 16 and the first bypass conduit 46 disposed immediately thereafter, It extends further radially outward beyond the radial span at the inlet end of the second bypass conduit 48 and is just below the inner surface of the third casing 44 surrounding the first fan, with a small radial spacing or Terminate with a gap.

  Correspondingly, the small diameter second fan 16 extends radially outward beyond both the inlet end of the core engine leading to the compressor 18 and the inlet end of the first bypass line 46; It ends with a small gap or gap inside the inlet end of the second casing 42 surrounding it.

  A row of fixed outlet guide vanes (OGV) 50 is coaxially disposed within the inlet end of the second bypass conduit 48 radially outward of the second fan 16 and aligned generally axially therewith. Yes. The OGV 50 has a suitable airfoil configuration for removing vortices of the compressed air 30 released from the radially outer tip portion of the first fan 14.

  The first fan 14 includes only a single stage or row of large first fan blades 52 that extend radially outward from the first support rotor disk 54. The second fan 16 is preferably a single stage or row of small second fan blades 56 that extend radially outward from the second support rotor disk 58 and are axially oriented. Then, it is arranged between the first fan blade 52 and the first bypass conduit 46.

  The first disk 54 is firmly connected to the first shaft 34, and the second disk 58 is firmly connected to the second shaft 36. In this manner, the small single-stage second fan 16 is immediately connected to the large single-stage first fan 14, and these fans are independently connected to the third turbine 26 and the second turbine 24. It is driven by this.

  The first fan blades 52 have a suitable airfoil configuration having a generally concave pressure side that is relatively smooth from the root to the tip of each blade and a generally convex suction side opposite the one, and any integral shape. It is characterized by the lack of the tip or the intermediate span side plate. In this way, the first fan blades can be made relatively thin to maximize the aerodynamic performance of the fan stage while reducing the centrifugal load that must be carried by the support rotor disk 54.

  Correspondingly, the smaller second fan blade 56 is also preferably configured with a generally concave pressure side extending smoothly from the root to the tip and a generally convex negative pressure side opposite to the generally concave pressure side. The second fan blade 56 is preferably configured in the airfoil cross-section to maximize its aerodynamic efficiency, and a correspondingly thin area reduces the weight and centrifugal loads carried by the support disk 58. Yes. Similar to the first blade 52, the second blade 56 is characterized by the absence of any integral tip or intermediate span side plate.

  Thus, the individual fan blades 52, 56 can be conventionally designed with their aerodynamic performance maximized while minimizing the weight and centrifugal loading of the blades carried by the corresponding disks 54, 58.

  Correspondingly, the high pressure compressor 18 can be conventionally configured to maximize its aerodynamic performance in a normal multi-stage axial compressor configuration. This arrangement includes several rows of vanes, which cooperate with corresponding rows of compressor blades commonly connected to the third shaft 38 to be driven by the HPT 22.

  The HPT 22 is a single-stage turbine including a stator nozzle at the outlet end of the annular combustor 20, and a single row of turbine blades extends radially outward from the support rotor disk. The compressor is driven by being firmly connected to the shaft 38.

  The IPT 24 is also a single stage turbine with corresponding stator nozzles cooperating with a single row of turbine blades, which turbine blades extend radially outward from the support rotor disk, the support rotor disk being The second fan 16 is driven by being firmly connected to the two shafts 36.

  The LPT 26 typically includes a number of stages with corresponding stator nozzles and cooperating rows of turbine blades that extend radially outward from the corresponding rotor disk; The rotor disk is firmly connected to the first shaft 34 and the first fan 14 on the upstream side is driven.

  The three concentric drive shafts or spools 34, 36, 38 are suitably mounted on several frames including a fan frame 60 at the front end of the engine and a rear frame 62 at the rear end of the engine. Yes. The frame is annular and includes a corresponding structural hub, which preferably supports various bearings and rotatably mounts three shafts within the engine for independent concentric rotation.

  The annular fan frame 60 is disposed between the second fan 16 and the compressor 18 in the axial direction and includes a row of fan struts 64 that extend from the central hub to the first bypass conduit 46. And the second bypass line 48 extend radially outwardly immediately after the OGV 50 near their inlet ends. The rear frame 62 includes a corresponding row of struts extending radially through the front end of the exhaust conduit 28 and supports the rear end of the drive shaft.

  The exhaust line 28 includes a main or first exhaust nozzle 66 connected to communicate with both the core engine and the first bypass line 46 and emits exhaust therefrom. The outlet end of the first bypass line 46 is arranged to communicate with the exhaust line 28, which receives the fan bypass flow from the first bypass line, and then this flow. Is mixed with the combustion gas discharged from the third turbine 26.

  The second bypass line 48 is arranged such that its outlet communicates with an auxiliary or second exhaust nozzle 68, which is preferably concentric with the first exhaust nozzle 66. In a preferred embodiment that reduces exhaust noise during engine operation, the outer bypass flow of air carried through the second bypass line 48 can then be selectively discharged from the second nozzle 68.

  The first exhaust nozzle 66 and the second exhaust nozzle 68 can have any conventional configuration and, in one embodiment, are conventionally configured so that the variable cross-sectional area is manipulated. The two nozzles have corresponding actuators that are preferably coupled to an engine controller 70 in the form of a digital computer, which controls the efficient operation of the engine during its operating cycle and flight envelope of the aircraft. The discharge flow rate range is suitably adjusted as necessary.

  In the exemplary embodiment shown in FIG. 1, the second exhaust nozzle 68 is concentrically disposed inside the first exhaust nozzle 66, and in an alternative embodiment, the second exhaust nozzle is outside the main nozzle 66. Can be arranged.

  For example, a conventional central cone or wedge 72 can be placed coaxially inside the rear end of the exhaust nozzle 28 to define the main exhaust nozzle 66. If it is desired to change the discharge flow rate range through the main nozzle 66, the wedge 72 can be suitably translated in the axial direction inside the exhaust line 28.

  In a normal tapered and divergent supersonic exhaust nozzle, the wedge 72 is diverged by increasing the diameter in the rearward direction to the peak of the maximum diameter, and the tapered conduit is symbolically A8 inside the exhaust conduit 28. It can be limited to terminate at the throat of the specified minimum flow range. The wedge then tapers backwards from the peak to reduce the diameter, limiting the divergent conduit terminating at the main nozzle, symbolically designated A9, with an increased outlet flow range.

  The auxiliary exhaust nozzle 68 can be suitably mounted inside the rear end of the main nozzle 66. Correspondingly, a row of hollow flow reversal struts 74 is arranged to communicate between the outlet end of the second bypass line 48 and the front or inlet end of the exhaust wedge 72.

  The reversing strut 74 provides a corresponding conduit that allows the outer bypass flow to be radially inward from the outer conduit 48 through the rear end of the inner conduit 46 and through the front end of the exhaust conduit 28 to the wedge 72. Carry. The second exhaust nozzle 68 is preferably positioned to communicate with the reversing post 74 inside the wedge 72 and selectively discharges an outer bypass flow from the exhaust wedge, if desired.

  In this configuration of the second exhaust nozzle 68, an annular ring valve is preferably mounted inside the wedge, and can be moved in parallel in the axial direction. If it is desired to discharge the compressed outer bypass flow from the main nozzle 66 to the main exhaust flow to reduce exhaust noise, the second nozzle 68 can be opened.

  Accordingly, the second nozzle 68 is configured as an acoustic nozzle, which injects compressed fan air into the main exhaust and mixes it, reducing its speed and attenuating noise. When control of noise attenuation or engine performance is no longer needed, the second nozzle 68 can be partially or completely closed appropriately.

  By introducing a variable cross-sectional area bypass injector (VABI) 76 at the rear end of the first bypass conduit 46, the performance of the variable cross-sectional area nozzles 66 and 68 can be supplemented. The VABI 76 can have any conventional configuration including a flap or ring valve to control the release of the inner bypass flow from the inner bypass line 46 to the common exhaust line 28.

  In this way, engine controller 70 controls and regulates the operation of variable exhaust nozzles 66, 68 and VABI 76, so that engine performance and efficiency can be maximized during its operating cycle. This includes independent and simultaneous control of the operating lines of the first and second fans 14 and 16 as they compress the airflow during operation. This includes the bypass ratio and stall margin associated with the double bypass fan.

  In the preferred embodiment shown in FIGS. 1 and 2, the first fan 14 and the second fan 16 have airfoil configurations with their respective fan blades 52, 56, and the third turbine 26 and the second turbine When power is given by 24, they are driven in reverse rotation. In particular, the corresponding generally concave pressure side of the fan blades 52, 56 faces in the opposite direction in the circumferential direction, so that the corresponding turbine 26, 24, which also has an airfoil configuration in the opposite direction in the circumferential direction. When driven by a moving blade, it rotates in reverse.

  The two stages of fan blades 52,56 operate in reverse rotation, improving aerodynamic performance and efficiency while eliminating external stages and components. For example, the large single-stage first fan 14 is directly connected so that the small single-stage second fan 16 communicates, and these fans are independently connected to their respective turbines 26 and 24 and rotated in reverse. .

  The row of fans OGV 50 communicates directly with the radially outer tips of the first fan blades 52 and removes compressed fan air vortices to the outer bypass conduit 48.

  The OGV 50 is directly followed by the fan strut 64 and the outer portion of the outer bypass conduit 48, and the second fan blade 56 is directly followed by the inner portion of the same strut in the inner bypass conduit 46.

  Thus, an inlet guide vane (IGV) is not necessary between the two fans 14, 16, and therefore, by eliminating the operating system specifically for it, the axial length of the engine is correspondingly reduced and the weight is reduced. And complexity can be reduced.

  Further, eliminating the corresponding OGV between the second fan 16 and the fan strut 64 can further reduce engine length and weight while increasing performance.

  The three spool independent variable double bypass turbofan engine shown in FIG. 1 can be used to benefit from the FLADE variable cycle engine without its mechanical complexity and limitations. Instead of using fan blades with FLADE side plates, the engine shown in FIG. 1 includes a large or extra large first stage fan 14 that is not integrated with FLADE. Thus, the first fan blade 52 can optimize aerodynamic efficiency and strength, and can have a relatively thin lateral area in which centrifugal loads are carried by the support rotor disk 54.

  The smaller second stage fan 56 is also configured without FLADE in it, so the second fan blade 56 is optimized for aerodynamic efficiency and strength, and the centrifugal load is effectively carried by the support rotor disk 58. It has a thin lateral area.

  The introduction of double bypass lines 46, 48 that cooperate with two fan stages will benefit from conventional FLADE designs without mechanical limitations. A large first stage fan blade 52 compresses incoming air 30 and its outer portion is carried directly through the OGV 50 to the outer bypass line 48. The inner portion of the compressed air from the first fan 14 is directly coupled to the second stage fan 16.

  The air is further compressed in the second fan blade 56, its radially outer portion is conveyed through the inner bypass line 46, and the central portion of the air from the second stage fan 16 is conveyed to the inlet of the high pressure compressor 18. .

  Both tip airflow flows from the two fans 14, 16 bypass the core engine through double bypass lines 46, 48. The inner bypass flow from the inner line 46 is discharged independently to the common exhaust line 28. The outer bypass flow in the outer conduit 48 is reversed from the rear strut 74 and is selectively emitted through the second acoustic nozzle 68 if desired to reduce engine noise.

  For use in supplying an acoustic nozzle 68 with a cross-sectional speed reversal, or for any other suitable purpose in a variable cycle engine, a double bypass turbofan engine without FLADE is relatively low pressure and low temperature. Create an outer bypass flow. The cold outer bypass flow can be used, for example, to cool the exhaust nozzle or provide a fluid shield around the exhaust flow.

  By eliminating FLADE from the engine, the aerodynamic penalties of the design with FLADE, flow path limitations, interstage leakage, and tip speed constraints associated with FLADE design associated with thicker airfoils are correspondingly Removed.

  Thus, two fan stages without FLADE can be optimized for airfoil design and higher fan efficiency, with specific fuel consumption correspondingly improved.

  Due to the counter-rotating configuration of the turbofan without FLADE, several rows of stator airfoils in the fan stage can be eliminated, thereby correspondingly reducing the length and weight of the engine. The use of reverse rotation can also eliminate the turbine nozzle between the second turbine 24 and the third turbine 26 operating in reverse rotation.

  In addition, the large first fan 14 is exposed to its own throttling region or back pressure in the outer bypass line 48 that terminates in the acoustic nozzle 68, thereby causing the first stage fan 14 and the first fan 14 to It is accordingly possible to control the operating lines of the two-stage fan 16 independently. The inner bypass line 46 is directly coupled to the second fan 16, and the bypass flow flows in parallel with the second exhaust nozzle 68 and discharges independently from the first exhaust nozzle 66.

  By independently controlling the fan stage, if desired, the variable cross-sectional area within the two nozzles 66, 68 can be controlled accordingly to optimize the performance of the fan in flight.

  Combining the benefits of a double bypass turbofan without the FLADE disclosed above offers the potential to significantly increase the area of supersonic transport aircraft or business jets over conventional variable cycle engines including designs with FLADE. . Preliminary engine cycle and aerodynamic analysis shows that the dual bypass turbofan engine disclosed above improves specific fuel consumption by 1% over the normal variable or adaptive cycle engine with FLADE configuration, Expects to reduce weight potentially by hundreds of pounds.

  Since the variable cycle engine specifically configured for propulsion of supersonic aircraft has a variety of configurations, the double bypass design without FLADE disclosed above is suitably adapted to complement the known conventional variable cycle engine. Can be corrected. The exhaust nozzle in a variable cycle engine has a variety of configurations that can be used to serve with the independent bypass flow from the outer and inner bypass lines that are directly coupled to the corresponding fan stages disclosed above.

  While what has been described herein, as well as exemplary embodiments of the invention, have been described, other modifications of the invention will become apparent to those skilled in the art from the teachings herein. Accordingly, it is desired that all modifications that come within the true spirit and scope of the invention be guaranteed in the appended claims.

  Accordingly, it is desired that the invention be protected by patent as defined and differentiated in the claims.

1 is a schematic axial view of a supersonic variable cycle turbofan aircraft engine. FIG. 2 is a plan view of a portion of the engine shown in FIG. 1 taken generally along line 2-2.

Explanation of symbols

DESCRIPTION OF SYMBOLS 10 Gas turbine engine 12 Center axis 14 1st fan 16 2nd fan 18 Compressor 20 Combustor 22 High pressure turbine (HPT)
24 Medium pressure turbine (IPT)
26 Low pressure turbine (LPT)
28 exhaust pipe 30 air 32 combustion gas 34 first drive shaft 36 second drive shaft 38 third drive shaft 40 first casing 42 second casing 44 third casing 46 first bypass conduit 48 second bypass conduit 50 outlet Guide wing (OGV)
52 First fan blade 54 First rotor disk 56 Second fan blade 58 Second rotor disk 60 Fan frame 62 Rear frame 64 Fan column 66 First exhaust nozzle 68 Second exhaust nozzle 70 Engine controller 72 Wedge 74 Inverted column 76 Variable Bypass Injector (VABI)

Claims (10)

In the variable cycle turbofan engine (10),
A large single-stage first fan (14) directly connected to communicate with a small single-stage second fan (16), wherein the first and second fans are independently turbines ( 26, 24) a first fan (14) connected to,
A first bypass line (46) surrounding a core engine including a compressor (18), a combustor (20), and a turbine (22) connected to the compressor (18), wherein the second fan A first bypass line (46) arranged to communicate with (16);
A second bypass line (48) surrounding the first bypass line (46), the second bypass line (48) being arranged to communicate with the first fan (14);
A first exhaust nozzle (66) coupled to communicate with both the core engine and the first bypass line (46);
An engine having a second exhaust nozzle (68) connected to communicate with the second bypass pipe (48).   A row of outlet guide vanes (50) arranged to communicate directly with the first fan (14) on the outer side of the second fan (16) within the inlet end of the second bypass pipe (48). The engine according to claim 1, further comprising:   The fan further includes a fan frame (60) disposed between the second fan (16) and the compressor (18), and the first and second bypass pipes are disposed behind the outlet guide vanes (50). The engine of claim 2 including a row of struts (64) extending radially outwardly through (46, 48).   The engine according to claim 3, wherein the first and second fans (14, 16) have diametrically opposite airfoil configurations and are counter-rotated when driven by the corresponding turbine (26, 24). A first casing (40) surrounding the compressor (18);
A second casing (42) that is spaced outward from the first casing (40) to limit the first bypass conduit (46);
An engine further comprising a third casing (44) that is spaced outward from the second casing (42) and defines the second bypass conduit (48);
The second fan (16) extends radially outward beyond the first bypass line (46) to the second casing (42);
The engine of claim 4, wherein the first fan (14) extends radially outwardly beyond the second bypass line (48) to the third casing (44). The first fan (14) is connected to a third turbine (26) by a first spool (34);
The second fan (16) is independently connected to the second turbine (24) in front of the third turbine (26) by a second spool (36);
The compressor (18) is independently connected to the first turbine (22) in front of the second turbine (24) by a third spool (38);
The first and second exhaust nozzles (66, 68) are disposed so as to communicate independently with the first and second bypass pipes (46, 48), respectively, and the first and second fans ( 14. The engine according to claim 5, wherein the operation lines 14 and 16) are independently controlled.   The engine according to claim 6, wherein the first and second exhaust nozzles (66, 68) are configured for variable cross-sectional areas.   The engine according to claim 6, wherein the second exhaust nozzle (68) is arranged concentrically with the inner side of the first exhaust nozzle (66). A central wedge (72) disposed coaxially with and inside the first exhaust nozzle (66);
An engine further comprising: a row of flow reversal struts (74) arranged to communicate between the second bypass line (48) and the wedge (72);
The second exhaust nozzle (68) is disposed so as to communicate with the reversing column (74) inside the wedge (72).
The engine according to claim 6.   The variable cross-section bypass injector (76) further disposed at a rear end of the first bypass line (46) for controlling a bypass flow to the first exhaust nozzle (66). engine.

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